Acoustically damped gas turbine engine

ABSTRACT

Disclosed is a gas turbine engine including a fan, a nacelle including a flutter damper forward of the fan, the flutter damper including an acoustic liner having a perforated radial inner face sheet and a radial outer back sheet, the acoustic liner configured for peak acoustical energy absorption at a frequency range that is greater than a frequency range associated with fan flutter, a chamber secured to the radial outer back sheet, the chamber in fluid communication with the acoustic liner, and the chamber configured for peak acoustical energy absorption at a frequency range associated with fan flutter modes, and the engine includes (i) the nacelle and a core cowl forming a convergent-divergent fan exit nozzle; (ii) a variable area fan nozzle capable of being in an opened and closed, the opened position having a larger fan exit area than the closed position; and/or (iii) the fan being shrouded.

BACKGROUND

Exemplary embodiments pertain to flutter dampers in gas turbinepropulsion systems and, more particularly, to flutter dampers in nacelleinlet structures.

Geared turbofan architectures, allow for high bypass ratio turbofans,enabling the use of low pressure ratio fans, which may be moresusceptible to fan flutter than high pressure ratio fans. Fan flutter isan aeromechanical instability detrimental to the life of a fan blade.

Accordingly, there is a need for a flutter damper which, by absorbingthe acoustic energy associated with the flutter structural mode, mayprevent the fan from fluttering, and which may be integrated into thereduced available space in an optimized propulsion system.

BRIEF DESCRIPTION

Disclosed is a gas turbine engine including: a fan; a nacelle includinga flutter damper disposed forward of the fan, the flutter damperincluding: an acoustic liner having a perforated radial inner face sheetand a radial outer back sheet, the acoustic liner being configured forpeak acoustical energy absorption at a frequency range that is greaterthan a frequency range associated with fan flutter; a chamber secured tothe radial outer back sheet, the chamber being in fluid communicationwith the acoustic liner, and the chamber being configured for peakacoustical energy absorption at a frequency range that is associatedwith one or more fan flutter modes; and wherein (i) the nacelle and acore cowl form a bypass duct, the bypass duct forming aconvergent-divergent fan exit nozzle; (ii) the gas turbine engineincludes a variable area fan nozzle capable of being in an openedposition and a closed position, wherein the opened position has a largerfan exit area than the closed position; and/or (iii) the fan is ashrouded fan.

Further disclosed is a method of reducing fan flutter in a gas turbineengine, including dampening acoustics with flutter damper disposed in anacelle, the flutter damper being forward of a fan, the flutter damperincluding an acoustic liner having a perforated radial inner face sheetand a radial outer back sheet, the acoustic liner being configured forpeak acoustical energy absorption at a frequency range that is greaterthan a frequency range associated with fan flutter, a chamber secured tothe radial outer back sheet, the chamber being in fluid communicationwith the acoustic liner, and the chamber being configured for peakacoustical energy absorption at a frequency range that is associatedwith one or more fan flutter modes, and wherein (i) the method includesdecreasing output pressure with a convergent-divergent fan exit nozzleformed in a bypass duct between a nacelle and a core cow; (ii) themethod includes decreasing output pressure with a variable area fannozzle in an opened position, wherein the opened position has a largerfan exit area than a closed position; and/or (iii) the fan being ashrouded fan.

Further disclosed is a method of reducing fan flutter in a gas turbineengine, including installing a flutter damper in a nacelle duct, theflutter damper being forward of the fan, the flutter damper including anacoustic liner having a perforated radial inner face sheet and a radialouter back sheet, the acoustic liner being configured for peakacoustical energy absorption at a frequency range that is greater than afrequency range associated with fan flutter, a chamber secured to theradial outer back sheet, the chamber being in fluid communication withthe acoustic liner, and the chamber being configured for peak acousticalenergy absorption at a frequency range that is associated with one ormore fan flutter modes, and applying a gas flow to the gas turbineengine, detecting fan flutter with the fan blades at a first anglerelative to inlet flow, and advancing a blade angle and determining forthe fan blades a second angle relative to inlet flow at which flutter isreduced, the second angle defining a closed angle for the fan bladesrelative to the first angle.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the fan operateswithin a flutter margin of between 2% and 10%. In addition to one ormore of the features described above, or as an alternative, furtherembodiments may include that the fan blades have a mean roughness ofless than about 28 Ra. In addition to one or more of the featuresdescribed above, or as an alternative, further embodiments may includethat the flutter damper has an impedance characteristic at one or moretarget frequencies defined as: f_(target)=f_(S,ND)+Ω·ND wherein f_(S,ND)is a resonance frequency corresponding to a structural mode of arotating component; ND is a nodal diameter count of the structural mode;and Ω is a rotational speed of the rotating component; and wherein theflutter damper has the following impedance characteristic at the one ormore targeted frequencies: R≥2ρc−3ρc≤X≤−0.6ρc wherein R is the real partof the impedance characteristic, X is the imaginary part of theimpedance characteristic, ρ is air density, and c is speed of sound.

BRIEF DESCRIPTION OF THE DRAWINGS

The following descriptions should not be considered limiting in any way.With reference to the accompanying drawings, like elements are numberedalike:

FIG. 1 is a schematic view of a gas turbine propulsion system;

FIG. 2 illustrates a perspective cross sectional view of a flutterdamper in a nacelle inlet;

FIG. 3 is a schematic view of a flutter damper in accordance with oneembodiment of the disclosure;

FIGS. 4A and 4B illustrate perspective views of one chamber of a flutterdamper in accordance with one embodiment of the disclosure;

FIG. 5 illustrates an array of chambers of flutter dampers integratedinto the nacelle inlet;

FIG. 6 is a perspective view of a portion of the nacelle inlet;

FIG. 7 illustrates a gas turbine engine according to an embodiment;

FIG. 8 illustrates a gas turbine engine according to an embodiment;

FIG. 9 illustrates a gas turbine engine according to an embodiment;

FIG. 10 illustrates a graph of certain engine design parameters;

FIG. 11 illustrates fan blades according to an embodiment;

FIG. 12 illustrates a method for mitigating fan flutter; and

FIG. 13 illustrates fan blades according to an embodiment.

DETAILED DESCRIPTION

A detailed description of one or more embodiments of the disclosedapparatus and method are presented herein by way of exemplification andnot limitation with reference to the Figures.

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct, while the compressor section 24 drives air along a coreflow path C for compression and communication into the combustor section26 then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viamultiple bearing systems 38. It should be understood that variousbearing systems 38 at various locations may alternatively oradditionally be provided, and the location of bearing systems 38 may bevaried as appropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. An engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The engine staticstructure 36 further supports bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis Awhich is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion. It will be appreciated that each of the positions of the fansection 22, compressor section 24, combustor section 26, turbine section28, and fan drive gear system 48 may be varied. For example, gear system48 may be located aft of combustor section 26 or even aft of turbinesection 28, and fan section 22 may be positioned forward or aft of thelocation of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and35,000 ft (10,688 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).

As illustrated in FIGS. 1 through 3, the engine 20 may include a nacelle100 with acoustic liner 101 at the radial inside of the nacelle inletskin 106. The acoustic liner 101 may have a perforated radial inner facesheet 108, i.e., facing a radial inside of a nacelle inlet 103,illustrated in FIG. 2, and a radial outer back sheet 110.

The acoustic liner 101 is designed to absorb energy that tends toproduce community noise. As such, for contemporary high bypass ratiopropulsion systems, the acoustic liner 101 typically provides for peakenergy absorption in the acoustic frequency range of about between 500and 2000 Hz, and is less effective outside this range. Fan flutter forsuch propulsion systems, however, typically occurs at a lower frequency,depending on the frequency and nodal diameter count of the criticalstructural mode. The structural frequency largely depends on the size ofthe fan, among other design parameters. Large fans tend to flutter atsmaller frequencies than small fans. Torsion modes tend to have higherfrequency than bending modes on any given fan, and either can becritical. The materials and construction techniques used to make the fanblades also have a significant influence on the frequency. Given therange of sizes, materials, and flutter critical modes in fans of moderngas turbine engines, the flutter frequency will typically occur at afrequency range of less than but not equal to 500 Hz, and morespecifically between 50 and 400 Hz, yet more specifically between 50 and300 Hz, and yet more specifically between 50 and 200 Hz.

In one embodiment, a flutter damper 102 is provided which may includethe acoustic liner 101 and a chamber 118 disposed radially exterior toand in acoustic communication with the acoustic liner 101. Also aflutter damper 102 without the acoustic liner 101 is considered part ofthe scope of this disclosure. As used herein, radially refers to theaxis A of the engine 20. Acoustic communication is provided through aperforation section 120 in the outer back sheet 110. In FIG. 2, theflutter damper 102 is illustrated as being disposed between a firstaxial forward nacelle bulkhead 114 and a second axial forward nacellebulkhead 116. The flutter damper 102, however, may be disposed anywherebetween a leading edge 111 of the fan 42 and a nacelle hilite 113, suchas flutter damper 102A disposed on the fan case 115 illustrated in FIG.1.

The flutter damper 102 may be configured to mitigate fan flutter byproviding peak energy absorption in the acoustic frequency rangeassociated with fan flutter modes, where such frequency range isreferred to herein as a flutter frequency range. The flutter damper mayhave desirable impedance characteristics at certain targeted flutterfrequencies, which may be defined as:f _(target) =f _(S,ND)+Ω·ND

In the equation above, the variable f_(S,ND) is the frequency, which ismeasured in units of Hertz, and which corresponds to a resonancefrequency of a structural mode of the fan blade, which typically may bea first or second bending mode with a certain nodal diameter count, ND.The variable ND is the nodal diameter count of the circumferentialpattern of the structural mode of the fan blade. The variable Ω is therotational speed of the fan, which is measured in the units ofrevolutions per second. The values for variable Ω may be chosen tocorrespond to conditions where fan flutter may typically occur, forexample, when the tip relative Mach number of the fan is between 0.85and 1.2 during standard-day, sea-level-static operation.

From the above equation, considering the nodal diameter constraints, thetargeted flutter frequency ranges may be defined to be:f _(S,ND)=frequency of first or second bending mode of fan with ND nodaldiameters1≤ND≤3Ω_(Mreltip=0.85)≤Ω≤Ω_(Mreltip=1.2)f _(target) =f _(S,ND)+Ω·ND

In the above equation, Mreltip is the tip relative Mach number for aradial outer tip of the fan blade, and the bending mode is a vibrationalmode of the fan blade. The symbol Ω_(Mreltip=0.85) denotes therotational speed where the tip relative Mach number is equal to 0.85;likewise, Ω_(Mreltip=1.2) denotes the rotational speed where the tiprelative Mach number is equal to 1.2, Of course, values greater orlesser than the aforementioned values are considered to be within thescope of the present disclosure.

Within the flutter frequency ranges associated with the first and secondbending mode, and more specifically at the targeted frequencies, theflutter damper may have the following impedance characteristics:R≥2ρc−3ρc≤X≤−0.6ρc

Again, these values may vary and fall within the scope of the presentdisclosure. The above equation references the impedance of the flutterdamper, defined as the complex ratio of the amplitude and phase ofpressure oscillations over the amplitude and phase of the acousticvelocity as a function of frequency. In addition, the equationreferences the real part of impedance is the resistance, which isvariable R, and the imaginary part of impedance is the reactance, whichis variable X. The variable ρ is the air density, and the variable c isthe sound speed, both being at the entrance to the flutter damper. Theresistance constraint on R may facilitate integration of the flutterdamper into acoustic liners, which typically have R values greater than2ρc in locations forward of the fan. The reactance constraint on Xoptimizes the flutter inhibiting capability of the device at operatingconditions typically encountered in commercial aircraft applications. Atcertain target frequencies, the flutter damper may satisfy the followingadditional constraint:

$0.0143 \leq \frac{{Vf}_{target}}{Sc} \leq 0.165$

Again, these values may vary and fall within the scope of the presentdisclosure. As illustrated in FIGS. 3, 4A and 4B, discussed in greaterdetail below, the chamber 118 has a width W, a height H, and a length L.In addition, the perforated section 120 disposed under the chamber 118has a width Wp and a length Lp, and the acoustic liner 101 has a heightH_(Li). Thus, in the above equation, the volume of the flutter damper102, which includes the volume (W×H×L) of chamber 118 and the volume(Wp×H_(Li)×Lp) of the acoustic liner 101 is variable V. The area of theperforated section 120 (Wp×Lp) disposed under the chamber 118 isvariable S. The units of V, S, c and f_(target) are chosen such that

$\frac{{Vf}_{target}}{Sc}$is non-dimensional.

Moreover, in one embodiment, a downstream edge of the chamber 118 may belocated at B/D≤0.35. In this equation, the variable B is the distancebetween the downstream edge of the chamber 118 and the fan tip leadingedge, and the variable D is the fan tip diameter at the leading edge ofthe fan blade.

Remaining with FIGS. 1-3, the illustrated flutter damper 102 designedaccording to the above constraints, has the benefit of being able to fitwithin smaller footprints of sized-optimized propulsion systems,providing a retrofittable solution to an existing engine inlet. Thus thedisclosed flutter damper 102 may help boost fan flutter margin withoutrequiring an inlet redesign. In addition, the flutter damper 102 mayprovide a relatively lightweight solution, that is, the low temperaturesof the inlet area may allow for the use of a metallic material,including aluminum, or a plastic or a composite, or a hybrid metallicand non-metallic material. Moreover, the flutter damper 102 may have ascalable design which can be oriented in an array of chambers, discussedin detail, below, and as illustrated in at least FIG. 5. For example,the array of chambers and may be placed around an engine inletcircumference to achieve a desired amount of flutter dampening volume.

As illustrated in FIGS. 4A and 4B, the perforation section 120 in theouter back sheet 110 may be rectangular in shape with length Lp andwidth Wp, where the length direction Lp corresponds to the engine axialdirection, and the width direction Wp corresponds to the enginecircumferential direction. For a contemporary high bypass ratiopropulsion system, which may have a fan diameter of about 80 inches, anda fan rotor hub-to-tip ratio of about 0.3, the length Lp may be aboutfour and half (4.5) inches for the chamber 118, and the width Wp may beabout twelve (12) inches for chamber 118. Each perforation section 120may have a hole-diameter of about thirty thousandths (0.030) of an inch.Of course, dimensions greater or lesser than the aforementioneddimensions, and non-rectangular shapes are considered to be within thescope of the present disclosure. This perforation geometry provides anopen area that may be about four and half (4.5) percent of the surfacearea (Lp×Wp) of the chamber 118 against the outer back sheet 110, whichmay be the same open area as a perforation section (not illustrated) inthe inner face sheet 108. Again, these dimensions may vary and remainwithin the scope of the present disclosure.

The chamber 118 may be sized to optimally dampen fan flutter at aspecific fan flutter frequency and nodal diameter. The nodal diametercount represents the nodal lines of vibrational modes observed for thefan blade, which typically may be between 1 and 3. The chamber 118 inFIG. 2, for example, is shaped as a rectangular box, and non-rectangularshapes are also within the scope of the disclosure, and may be sizedbased on an observed flutter frequencies and nodal diameters for a givenengine. For example, if an engine has an observable flutter mode at afrequency of about 150 Hz with nodal diameter 2, the chamber 118 may besized according to that flutter mode and nodal diameter.

The box shape, as illustrated in FIG. 4A, may have a top surface 122roughly defined by a width-length (W×L) area, where the length directionL corresponds to the engine axial direction, and the width direction Wcorresponds to the engine circumferential direction. The box shape mayalso have a front surface 124 and a back surface 125, each roughlydefined by a height-width (H×W) area, where the height direction H forthe chamber 118 may correspond to an engine radial direction. The boxshape may further have a side surface 126 roughly defined by aheight-length (H×L) area. Again, these dimensions may vary and remainwithin the scope of the present disclosure.

For the exemplary embodiment, the chamber 118 is twelve (12) incheswide, as referenced above, and the chamber width-height-length (W×H×L)volume may be three hundred twenty four (324) cubic inches, and theheight H may be equal to, or less than, six (6) inches.

Turning now to FIGS. 4A and 4B, the box shaped chamber 118 may have abottom edge 128 that geometrically conforms to the annular and axialprofile shape of the nacelle inlet 103. Extending axially andcircumferentially outwardly from the bottom edge 128 of the chamber 118is a mounting flange 130 for affixing the chamber 118 to an existingnacelle inlet 103. As such, the bottom face 131 of the chamber 118 maybe formed by the radial outer back sheet 110 of the acoustic liner 101.

The chamber 118 may also include first and second stiffening structures132, 134. The stiffening structures 132, 134 may have a substantially“C” shape, when viewing into the side surface 126 of the chamber 118,which protrudes outwardly from the top 122, front 124 and back 125surfaces of the chamber 118. The stiffening structures 132, 134 maydivide the top surface 122 of the chamber 118 in substantially equalportions in the width direction W. The stiffening structures 132, 134may tune the structural resonance frequencies of the chamber 118 awayfrom the fan flutter frequencies to avoid fan flutter inducing resonancein the chamber 118. For example, the stiffening structures 132, 134 maytune the structural resonance frequencies of the relatively large, flattop surface 122 of the chamber 118 out of the targeted flutter frequencyrange. In addition, the stiffening structures 132, 134 add structuralrigidity and may allow for a lightweight design of the chamber 118.

One or more weep holes 136 may be provided to allow for water or fluidegress. The placement of the weep holes 136 is selected to be below theengine centerline.

Turning now to FIGS. 5 and 6 a circumferential array 138 of chambers118, including fourteen (14) chambers 118, is disposed about the nacelleinlet 103, with each of the chambers 118 having a perforated section.Disposing the chambers 118 in this type of circumferential array 138achieves a desired damping volume.

Turning now to FIGS. 7-9, three embodiments of a gas turbine engine 202,204, and 206 are illustrated. Each of the three embodiments may includea nacelle 208 fitted with an acoustic damper 210. The acoustic damper210 may be located axially between a hilite 212 disposed at an engineinlet 213 and a fan 214. As above, the acoustic damper 210 may have anacoustic liner 216 located against an inlet-flow facing surface 218. Theacoustic damper 210 may also include a resonator chamber 220 locatedradially outside of the liner 216.

Radially within the nacelle 208, the engines 202, 204, 206 may includean engine core 222 surrounded by an engine core cowl 224. The enginecore 222 may include a compressor module 226, a combustor 228 and aturbine module 230. A core bypass area 232 may be located radiallybetween the cowl 222 and the flow facing surface 218 of the nacelle 208.The bypass area 232 in the illustrated embodiments may be a high-bypassarea.

FIG. 10 illustrates a graph of thrust on the abscissa 250 and a ratio ofoutput pressure to input pressure on the ordinate 252 for a constantoutlet area gas turbine engine. As illustrated in FIG. 10, outputpressure may increase as thrust increases for normal fan operationconditions. The bottom curve 254 is a fan operating curve for a thrusttarget range. For a given thrust target, the top curve 256 illustratesthe occurrence of fan flutter, i.e., blade radial outer tips bending inthe axial forward direction. Fan flutter may occur if output pressureincreases over input pressure by small margin 258 over the fan'soperating parameters.

Flutter may be realized during intervals of high engine thrust, e.g.,during takeoff and climb. In addition, the flutter margin 258 maydecrease as the fan blade becomes impacted by increased leading edgeroughness, decreased blade cleanness, and compromised blade clearances,all of which may occur during normal use. The margin for a blade iscloser to optimal when the blade is newly manufactured. For commercialaircraft, it is not uncommon to require a newly manufactured orrefinished surface texture mean roughness of all airfoil surfaces to beless than 28 Ra (Roughness Average). Additional factors which may impactthe flatter margin include engine thrust deflections, cross winds, andother aerodynamic and mechanical factors.

As illustrated in FIG. 10, reducing output pressure, e.g., duringtakeoff, may reduce the likelihood of fan flutter during that flightphase. Output pressure may be reduced by increasing output area 302, asillustrated in FIG. 7 with a variable area fan nozzle 304. A variablearea fan nozzle 304 may pivot and/or move axially rearward to increasethe output area. Another option to reduce output pressure, asillustrated in FIG. 8, may be to permanently fix the variable area fannozzle 304 in semi-opened state. Yet another option to reduce outputpressure, as illustrated in FIG. 9, may be to design the nacelle 208and/or engine cowl 224 so that the engine bypass 232, downstream of thefan 214, forms a convergent-divergent nozzle. That is, a first bypassarea A1, e.g., axially at the compressor module 226, may be larger thana second bypass area A2, e.g., axially at the combustor module 228. Athird bypass area A3, e.g., axially at the turbine module 230 may alsobe larger than the second bypass area A2. Mathematically, this arearelationship may be represented as A1>A2; A2<A3.

FIG. 11 illustrates another option which may lower output pressure. Twoblades 350, 352 of the fan are illustrated. At the leading edge 354 andtrailing edge 356 of one of the blades 352, force triangles are providedillustrating absolute velocities V1, V2, blade velocities U1, U2,relative velocities C1, C2, axial flow velocities C1AX, C2AX and swirlvelocities C1U, C2U. Closing the blades 350, 352 means that, relative toa more efficient configuration, the blades 350, 352 are turned aroundeach blade radial center 358, 360, pursuant to arrows 362, 364. Theresulting origination of the blades 350, 352 increases the amount ofblade surface area facing the nacelle inlet flow, which decreases flowvelocities, C1AX, C2AX, and may lower output pressure.

A determination of the closing angle can be made by installing flutterdampers into an engine known to experience fan flutter, and wind tunneltesting the engine until flutter ceases. The close angle obtained withthe installation of the flutter damper will be smaller than without suchinstallation, and the engine will run more efficiently. For example, asillustrated in FIG. 12, flutter dampers are installed into an engine inSTEP 302. Then sensors are attached to detect fan flutter with theengine set at high thrust to simulate takeoff at STEP 204. When fanflutter is detected, the blade angle is advanced to a closed position,e.g., one degree per advancement. Once flutter has ceased, the angle isrecorded. This angle is used to manufacture blades that can withstandflutter in high thrust conditions. The close angle will be less thanwith a similarly tested engine in which the flutter damper is notinstalled. As a result, the engine will run more efficiently.

FIG. 13 illustrates another option which may lower output pressure. Fanblades 400, e.g., blade 402, may be connected in part via a radialmid-span shroud 404. The shroud 404 may enable the use of more blades400 having a shorter cord length than an unshrouded blade configuration.The shorter cord length in the shrouded blade provides less skinfriction drag per blade than the unshrouded blade. The inherently loweroutput pressure of the shrouded blade would provide a larger fluttermargin. Thus, the installation of the flutter damper would furtherprolong the onset of flutter, prolong periods of time between bladesurface refurbishing, and lengthen the useful life of a blade.

Each of the solutions in FIGS. 7-13 provide a solution which reduces thelikelihood of fan flutter without making more robust mechanical changesto the engine that are otherwise required to reduce output pressure. Asa result, the efficiency of the engines 202, 204, 206 may increase,which may reduce fuel consumption, community noise, and engine wear.

The term “about” is intended to include the degree of error associatedwith measurement of the particular quantity based upon the equipmentavailable at the time of filing the application. For example, “about”can include a range of ±8% or 5%, or 2% of a given value.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the presentdisclosure. As used herein, the singular forms “a”, “an” and “the” areintended to include the plural forms as well, unless the context clearlyindicates otherwise. It will be further understood that the terms“comprises” and/or “comprising,” when used in this specification,specify the presence of stated features, integers, steps, operations,elements, and/or components, but do not preclude the presence oraddition of one or more other features, integers, steps, operations,element components, and/or groups thereof.

While the present disclosure has been described with reference to anexemplary embodiment or embodiments, it will be understood by thoseskilled in the art that various changes may be made and equivalents maybe substituted for elements thereof without departing from the scope ofthe present disclosure. In addition, many modifications may be made toadapt a particular situation or material to the teachings of the presentdisclosure without departing from the essential scope thereof.Therefore, it is intended that the present disclosure not be limited tothe particular embodiment disclosed as the best mode contemplated forcarrying out this present disclosure, but that the present disclosurewill include all embodiments falling within the scope of the claims.

What is claimed is:
 1. A gas turbine engine comprising: a fan; a nacelleincluding a flutter damper disposed forward of the fan, the flutterdamper comprising: an acoustic liner having a perforated radial innerface sheet and a radial outer back sheet, wherein a perforated sectionis defined by perforations in the radial outer back sheet, the acousticliner being configured for peak acoustical energy absorption at afrequency range that is greater than a frequency range associated withfan flutter; a circumferential array of chambers secured around acircumference of a nacelle inlet, the chambers being circumferentiallyspaced from one another and cumulatively achieving a desired dampeningvolume, each of the chambers being secured to the radial outer backsheet, and being in fluid communication with the acoustic liner, each ofthe chambers configured so that the perforated section is disposedthereunder and has a width W_(p), a length L_(p) and a heightcorresponding to a height of the acoustic liner H_(Li), wherein L_(p) isdefined along an engine axial direction, and each of the chambersdefining a rectangular box with a front surface, a back surface,opposing side surfaces, and a bottom edge extending axially andcircumferentially outwardly from each of the surfaces to define amounting flange that geometrically conforms to an annular and axialprofile of the nacelle inlet and that mounts each of the chambers to thenacelle inlet, whereby a bottom surface of each of the chambers isformed by the radial outer back sheet, and wherein each of each of thechambers has a width W, a length L, and a height H, wherein L is definedalong the engine axial direction, wherein an acoustic volume V of theflutter damper is a sum of W_(p)×H_(Li)×L_(p) and W×H×L, and wherein theperforated section and each of the chambers is configured so thatL>L_(p) within each of the chambers, and each of the chambers isconfigured for peak acoustical energy absorption at a frequency rangethat is associated with one or more fan flutter modes; and wherein: (i)the nacelle and a core cowl form a bypass duct, the bypass duct forminga convergent-divergent fan exit nozzle; and/or (ii) the gas turbineengine includes a variable area fan nozzle, the variable area fan nozzleis capable of being in an opened position and a closed position, whereinthe opened position has a larger fan exit area than the closed position;and/or (iii) the fan is a shrouded fan.
 2. The gas turbine engine ofclaim 1, wherein the fan operates within a flutter margin of between 2%and 10%.
 3. The gas turbine engine of claim 2, wherein blades of the fanhave a mean roughness of less than about 28 Ra.
 4. The gas turbineengine of claim 1, including a variable area fan nozzle, and wherein thevariable area fan nozzle is fixed in a semi-opened state.
 5. The gasturbine engine of claim 4, wherein the fan operates within a fluttermargin of between 2% and 10%.
 6. The gas turbine engine of claim 5,wherein blades of the fan have a mean roughness of less than about 28Ra.
 7. The gas turbine engine of claim 4, wherein: the flutter damperhas an impedance characteristic at one or more target frequenciesdefined as:f _(target) =f _(S,ND)+Ω·ND wherein f_(S,ND) is a resonance frequencycorresponding to a structural mode of a rotating component; ND is anodal diameter count of the structural mode; and Ω is a rotational speedof the rotating component; and wherein the flutter damper has thefollowing impedance characteristic at the one or more targetedfrequencies:R≥2ρc−3ρc≤X≤−0.6ρc wherein R is the real part of the impedancecharacteristic, X is the imaginary part of the impedance characteristic,ρ is air density, and c is speed of sound.
 8. The gas turbine engine ofclaim 1, wherein: the flutter damper has an impedance characteristic atone or more target frequencies defined as:f _(target) =f _(S,ND)+Ω·ND wherein f_(S,ND) is a resonance frequencycorresponding to a structural mode of a rotating component; ND is anodal diameter count of the structural mode; and Ω is a rotational speedof the rotating component; and wherein the flutter damper has thefollowing impedance characteristic at the one or more targetedfrequencies:R≥2ρc−3ρc≤X≤−0.6ρc wherein R is the real part of the impedancecharacteristic, X is the imaginary part of the impedance characteristic,ρ is air density, and c is speed of sound.